1. Field of the Invention
This invention pertains to a gas turbine airfoil and in particular to a cooling construction for its leading edge.
2. Brief Description of the Related Art
Airfoils of gas turbines, turbine rotor blades, and stator vanes, require extensive cooling in order to keep the metal temperature below a certain allowable level and prevent damage due to overheating. Typically such airfoils are designed with hollow spaces and a plurality of passages and cavities for cooling fluid to flow through. The cooling fluid is typically air bled from the compressor having a higher pressure and lower temperature compared to the gas travelling through the turbine. The higher pressure forces the air through the cavities and passages as it transports the heat away from the airfoil walls. The cooling construction further usually includes film cooling holes leading from the hollow spaces within the airfoil to the external surfaces of the leading and trailing edge as well as to the suction and pressure sidewalls.
The leading edge of a turbine blade is one of the areas that faces the hottest gas flow conditions, and is thus one of the most critical areas to be cooled. It also has the particularity to have a strong surface curvature and thus a highly accelerated flow from each side of the stagnation line. For very hot gas temperature conditions, cooling the leading edge with an internal cooling passage is usually not sufficient, requiring additional rows of holes drilled into the leading edge to pick-up some heat directly through the holes and to provide a layer of coolant film on the external surface. However the interaction of the coolant flow ejected from theses rows of holes and the main hot gas flow can be difficult to predict, especially in situations where the stagnation line position can be uncertain due to changes of incidence angles. For this reason extensive studies have been performed on several leading edge film cooling configurations, including cylinders and blunt body models that simulate the leading edge of a turbine airfoil.
In the state of the art generally, the film cooling holes extending from cooling passages within the airfoil to the leading edge are positioned at a large angle to the leading edge surface and are designed with a small length to diameter ratio. Typically, the angle between the cooling hole axis and the leading edge surface is significantly greater than 20 degrees and the ratio of the cooling hole length to the cooling hole diameter is about 10, typically less than 15. Such holes are drilled by an electro-discharge machining process and, more recently, by a laser drilling process. Such film cooling holes provide good convective cooling of the leading edge of the airfoil due to the cumulative convective cooling area of all the film cooling holes together that are positioned between the root and the tip of the airfoil leading edge. The cooling air that exits the film cooling holes provides further cooling by means of a film that passes along the surface of the airfoil leading edge.
The establishment of a cooling film by a number of exit holes along the leading edge is sensitive to the pressure difference across the exit holes. While too small a pressure difference can result in an ingestion of hot gas into the film cooling hole, too large a pressure difference can result in the cooling air blowing out of the hole and will not reattach to the surface of the airfoil for film formation.
Furthermore, the short length-to-diameter ratio of the film cooling holes and the large angle between the hole axes and the leading edge surface can lead to the formation of vortices about the exit holes. This results in a high penetration of the cooling film away from the surface of the airfoil and in a decrease of the film cooling effectiveness about the leading edge of the airfoil.
One way to provide better film cooling of the airfoil surface is to orient the film cooling holes at a shallower angle with respect to the leading edge surface. This would decrease the tendency of vortex formation. However, a more shallow angle results in a larger length to diameter ratio of the film cooling hole, which exceeds the capabilities of today's laser drilling machines.
EP 0 924 384 discloses an airfoil with a cooling construction of the leading edge of an airfoil that provides improved film cooling of the surface. The disclosed airfoil includes a trench that extends along the leading edge and from the root to the tip of the airfoil. The apertures of the film cooling holes are positioned within this trench in a continuous straight row. The cooling air bleeds to both sides of these apertures and provides a uniform cooling film downstream and to both sides of the airfoil.
U.S. Pat. No. 5,779,437 provides a cooling system for the showerhead region, in which there is a multitude of passages, wherein each passage has a radial component and a downstream component relative to the leading edge axis, and the outlet of each passage has a diffuser area formed by conical machining, wherein the diffuser area is recessed in the wall portion downstream of the passage.
EP 1 645 721 discloses an airfoil having several film cooling holes at the leading edge with exit ports. The film cooling holes have a sidewall that is diffused in the direction of the tip of the airfoil at least over a part of the film cooling hole. Furthermore, the film cooling holes each have flare-like contour near the outer surface of the leading edge. The film cooling holes are stated to provide an improved film cooling effectiveness due to reduced formation of vortices and decreased penetration depth of the cooling air film.